Complex cycles

ABSTRACT

A complex cycle gas turbine engine ( 103 ) for an aircraft with hydrogen fuel supply. The gas turbine engine comprises, in fluid flow series, a core gas turbine ( 105 ) and a recuperator ( 701 ) for heating hydrogen fuel prior to combustion by turbine exhaust products.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority from United Kingdom Patent ApplicationNos 20 15 452.2, 20 15 453.0, and 20 15 454.8, all filed Sep. 30, 2020,and United Kingdom Patent Application No 20 17 014.8 filed Oct. 27,2020, each of which is incorporated herein by reference in its entirety.

TECHNICAL FIELD

This disclosure relates to complex cycle hydrogen-fuelled aero gasturbine engines.

BACKGROUND

In order to limit emissions of carbon dioxide, use of hydrogen as analternative to hydrocarbon fuel in gas turbine engines has historicallyonly been practical in land-based installations. Such engines aretypically supplied with hydrogen derived from natural gas via concurrentsteam methane reformation, which hydrogen is injected into large-volumeseries staged dry low NO_(x) burners. This type of burner is notsuitable for use in an aero engine primarily due to its size and thedifficulties in maintaining stable operation during transientmanoeuvres.

Experimental programmes have been conducted to develop aero enginesoperable to be fuelled with hydrogen, however these have typically beenhigh-Mach afterburning turbojets or expander cycles and thus notpractical for use on civil airliners operating in the Mach 0.8 to 0.85regime.

There is therefore a need for technologies for combustion of hydrogen inaero gas turbine installations, in particular around the overall enginecycle to for example minimise fuel consumption, the fuel delivery systemto for example meter fuel accurately, and the fuel injection system tofor example minimise emissions.

SUMMARY

The invention is directed towards complex cycle gas turbine engines foraircraft with hydrogen fuel supply. In an aspect, one such enginecomprises, in fluid flow series, a core gas turbine and a recuperatorfor heating hydrogen fuel prior to combustion by turbine exhaustproducts.

In an embodiment, the core gas turbine engine comprises:

a low-pressure spool having a low-pressure compressor driven by alow-pressure turbine;

a high-pressure spool having a high-pressure compressor driven by a lowhigh-pressure turbine.

In an embodiment, the engine further comprises a fan driven by thelow-pressure spool, optionally via a reduction gearbox.

In an embodiment, the core gas turbine comprises a combustor between thehigh-pressure compressor and the high-pressure turbine, the combustorhaving a liner cooled in a regenerative configuration by the hydrogenfuel.

In an embodiment, the engine further comprises a fuel turbine forreceiving heated hydrogen fuel from the recuperator to drive a load.

In an embodiment, the load is:

an electrical generator;

a fuel pump for pumping the hydrogen fuel; and/or

a low-pressure spool or a high-pressure spool in the core gas turbine.

In an embodiment, the engine further comprises an intercooler betweenthe low-pressure compressor and the high-pressure compressor for coolinglow-pressure compressor discharge air by the hydrogen fuel.

In an embodiment, the engine further comprises a second recuperator forfurther heating hydrogen fuel received from the intercooler.

In an embodiment, the second recuperator is stationed downstream of therecuperator.

In an embodiment, the engine further comprises a reheat combustor.

In an embodiment, the reheat combustor is stationed between thehigh-pressure turbine and the low-pressure turbine.

In an embodiment, the core gas turbine engine comprises a multi-stageturbine and the reheat combustor is stationed between two of the stageson the turbine.

In an embodiment, the engine further comprises a fuel delivery systemfor delivering the hydrogen fuel from a cryogenic storage system to therecuperator, the fuel delivery system including a pump, a meteringdevice, and a fuel heating system for heating the hydrogen fuel.

In an embodiment, the fuel heating system comprises a vaporiserconfigured to vaporise liquid hydrogen from the cryogenic storagesystem.

In an embodiment, the vaporiser comprises a fuel offtake for diverting aportion of the hydrogen fuel from a fuel conduit for combustion in aburner located in heat exchange relationship with the fuel conduit.

In an embodiment, the vaporiser comprises a boil volume or an electricheating element for initial heating of liquid hydrogen if no vaporisedhydrogen fuel is available.

In an embodiment, the metering device is a fixed orifice and flow rateis controlled by varying the pressure ratio across the orifice.

In an embodiment, the metering device comprises a sonic fixed orificeconfigured to operate in a choked condition, and flow rate is controlledby varying pressure upstream of the sonic fixed orifice.

In an embodiment, the heating system comprises a heater for heating ofthe hydrogen fuel following metering by the metering device.

In an embodiment, the heater comprises a fuel offtake for diverting aportion of the hydrogen fuel from a fuel conduit for combustion in aburner located in heat exchange relationship with the fuel conduit.

In an embodiment, the heating system comprises one or more heatexchangers for heating the hydrogen fuel by heat from the core gasturbine.

In an embodiment, the one or more heat exchangers are oil-fuel heatexchangers for cooling engine oil or gearbox oil by the hydrogen fuel.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only with referenceto the accompanying drawings, which are purely schematic and not toscale, and in which:

FIG. 1 shows a hydrogen-fuelled airliner comprising hydrogen-fuelledturbofan engines;

FIG. 2 is a block diagram identifying the flow of hydrogen fuel;

FIG. 3 shows a fuel delivery system;

FIG. 4 shows a fuel-oil heat exchange arrangement for the fuel heatingsystem of the fuel delivery system of FIG. 3;

FIG. 5 shows a vaporiser of the fuel delivery system of FIG. 3;

FIG. 6 shows a metering device of the fuel delivery system of FIG. 3;

FIG. 7 shows a complex cycle including a recuperator;

FIG. 8 shows another complex cycle including a recuperator and a fuelturbine;

FIG. 9 shows another complex cycle including a recuperator andrecuperative combustor cooling including a recuperator;

FIG. 10 shows another complex cycle including an intercooler andtwin-pass recuperation;

FIG. 11 shows another complex cycle including an intercooler,recuperation and inter-turbine reheat;

FIGS. 12A and 12B show two possible arrangements of the fuel injectionsystem of the engines of FIG. 1;

FIGS. 13A and 13B show two embodiments of a rim injector block;

FIG. 14A shows the one configuration of the rim injector in crosssection and FIG. 14B shows simulated equivalence ratios downstream ofthe rim injector of FIG. 14A;

FIG. 15A shows the one configuration of the rim injector in crosssection and FIG. 15B shows simulated equivalence ratios downstream ofthe rim injector of FIG. 15A;

FIG. 16 shows an embodiment of a converging jet injector block;

FIG. 17A shows the converging jet injector block in cross section andFIG. 17B shows simulated equivalence ratios downstream of the convergingjet injector block of FIG. 17A;

FIGS. 18A and 18B show two embodiments of a jet matrix injector block;

FIG. 19A shows the jet matrix injector block in cross section and

FIG. 19B shows simulated equivalence ratios downstream of the jet matrixinjector block of FIG. 19A; and

FIGS. 20A and 20B show the invariance of injector block size with powerscaling of engine and fuel injection system.

DETAILED DESCRIPTION

A hydrogen-fuelled airliner is illustrated in FIG. 1. In this example,the airliner 101 is of substantially conventional tube-and-wing twinjetconfiguration with a central fuselage 102 and substantially identicalunderwing-mounted turbofan engines 103.

In the present embodiment, the turbofan engines 103 are geared turbofanengines. A hydrogen storage tank 104 located in the fuselage 104 forhydrogen fuel supply is connected with core gas turbines 105 in theturbofan engines 103 via a fuel delivery system. In the presentembodiment, the hydrogen storage tank 104 is a cryogenic hydrogenstorage tank and thus stores the hydrogen fuel in a liquid state, in aspecific example at 20 kelvin. In this example, the hydrogen fuel ispressurised to a pressure from around 1 to around 3 bar, in a specificexample 2 bar.

A block diagram identifying the flow of hydrogen fuel is shown in FIG.2.

Hydrogen fuel is obtained from the hydrogen storage tank 104 by the fueldelivery system 201 and supplied to each core gas turbine 105. In theFigure, only one of the core gas turbines is shown for clarity. In thisillustrated embodiment, the core gas turbine 105 is a simple cycle gasturbine engine. In other embodiments, as will be described withreference to FIGS. 7 to 11, complex cycles may be implemented viafuel-cooling of the gas path.

Referring again to FIG. 2, the core gas turbine 105 comprises, in fluidflow series, a low-pressure compressor 202, an interstage duct 203, ahigh-pressure compressor 204, a diffuser 205, a fuel injection system206, a combustor 207, a high-pressure turbine 208, a low-pressureturbine 209, and a core nozzle 210. The high-pressure compressor 204 isdriven by the high-pressure turbine 208 via a first shaft 211, and thelow-pressure compressor 203 is driven by the low-pressure turbine 209via a second shaft 212. It will be appreciated that in alternativeembodiments, the core gas turbine could be of three-shaft configuration.

As will be described further with reference to FIG. 12A onward, the fuelinjection system 206 may be a direct fuel injection system.

As described previously, in the present embodiment, the turbofan engines103 are geared turbofan engines. Thus in operation the low-pressureturbine 209 drives a fan 213 via a reduction gearbox 214. The reductiongearbox receives input drive from the second shaft 212 and providesoutput drive to the fan 213 via a fan shaft 215. In an embodiment, thereduction gearbox 214 is an epicyclic reduction gearbox. In a specificembodiment, it is a planetary reduction gearbox. Alternatively, it maybe a star reduction gearbox, or a compound epicyclic reduction gearbox.As a further alternative, the reduction gearbox 214 could be alayshaft-type reduction gearbox or any other type of reduction gearbox.It will also be appreciated that the principles disclosed herein may beapplied to a direct-drive type turbofan engine, i.e. in which there isno reduction gearbox between the low-pressure turbine and the fan.

Fuel Delivery System

In operation, the fuel delivery system 201 is configured to obtainhydrogen fuel from the hydrogen storage tank 104 and provide it to thefuel injection system 206 in the core gas turbine 105. FIG. 3 is a blockdiagram illustrating the fuel delivery system 201 in greater detail.

The fuel delivery system 201 comprises a pump 301, a metering device302, and a fuel heating system for heating the hydrogen fuel to aninjection temperature for the fuel injection system 206. In anembodiment, a vent system (not shown) may be included in the fueldelivery system 201 close to the fuel injection system 206 to venthydrogen fuel should a rapid shut-off be required, for example inresponse to a shaft-break event. It is envisaged that the vent systemmay vent the excess hydrogen fuel into the bypass duct of the turbofanengine 103, or alternatively vent it outside of the nacelle of theengine 103. An igniter may be provided to flare off the excess hydrogenin a controlled manner.

In the present embodiment, the pump 301 is high-speed centrifugal pump.In a specific embodiment, it is configured to operate at 50000 rpm ormore. In a specific embodiment, the centrifugal pump comprises an axialinducer to minimise the required inlet pressure and to accommodatemultiphase flow in addition to the centrifugal impeller for developingthe majority of the required pressure rise. In an alternativeembodiment, a piston-type pump could be used.

In an embodiment, the pump 301 is located in the hydrogen storage tank104. In this way leakage of hydrogen fuel past pump seals etc. isaccommodated.

In an embodiment, the pump 301 is driven by a fuel turbine, as will bedescribed with reference to FIG. 8.

Alternatively, the pump 301 could be driven by an air turbine suppliedwith compressor bleed, for example bleed from the high-pressurecompressor 204. Alternatively, combustion products from the combustor207 may be used to drive a dedicated turbine for driving the pump 301.In another embodiment, the pump 301 is driven via an electrical machine.In an embodiment, the drive means for the pump 301 are also located inthe hydrogen storage tank 104.

In this embodiment, the metering device 302 is configured to meter therequired quantity of fuel for the current fuel demand of the core gasturbine 105.

As will be appreciated, it is desirable to increase the temperature ofthe fuel from the 20 kelvin cryogenic storage condition to a temperaturemuch closer to the firing temperature of the core gas turbine; of coursethis is subject to the constraint of not exceeding the autoignitiontemperature of the hydrogen fuel prior to admission into the combustor207. In an example, the injection temperature is from 250 to 300 kelvin,for example 280 kelvin.

In the present embodiment, the fuel heating system comprises a vaporiser303 for heating of the hydrogen fuel to implement a phase change. In thepresent embodiment, this takes place between the pump 301 and themetering device 302. In this way the metering device 302 meters gaseoushydrogen fuel. It is contemplated that in other embodiments, the orderof operations may change.

In an embodiment, the vaporiser 303 is configured to raise thetemperature of the hydrogen fuel to the required injection temperature.Thus, in such a configuration, the metering device 302 meters thehydrogen fuel at the injection temperature.

In another embodiment, the vaporiser 303 is configured to raise thetemperature of the hydrogen fuel to a metering temperature less than theinjection temperature. This could for example be from 100 to 200 kelvin,for example 150 kelvin. This reduces the risk of damage to electronicdevices used for sensing temperature, pressure etc.

Further heating is implemented following the metering of hydrogen fuelby the metering device 302. In this example, this is achieved with aheater 304. The configuration of the vaporiser 303 and heater 304 may besubstantially similar, and an example will be described further withreference to FIG. 10.

Additionally or alternatively, the fuel heating system may comprise oneor more heat exchangers for raising the temperature of the hydrogen fuelby use of rejected heat from the core gas turbine 105. As will bedescribed further with reference to FIGS. 7 to 11, this may be achievedby implementing a complex cycle configuration, for example fuelrecuperation, intercooling, etc.

However, even in a simple cycle configuration as contemplated herein,this fuel heating may be achieved by, for example, cooling one or moreof the various oil systems in the core gas turbine 105. A specificexample of such a configuration is shown in FIG. 4, in which the fuelheating system comprises a fuel-oil heat exchanger 401 for coolinglubricating oil from the reduction gearbox 214. In an example, even witha 99 percent efficient gearset, at maximum thrust it may still berequired to reject around 750 kilowatts of heat from the gearbox oilsystem, which represents a significant opportunity for raising thetemperature of the hydrogen fuel. It will be appreciated that otherengine oil, such as main bearing lubrication oil, may also be cooled ina similar manner. It will also be appreciated that cooling air systemsmay be cooled in a similar manner, with high-pressure compressor 204discharge air being cooled by heat exchange with the hydrogen fuel priorto being delivered to the high-pressure turbine 208 for cooling thereof.

In a simple cycle configuration it has been determined that due to thesignificant heat capacity of the hydrogen fuel, even if it is utilisedas a heatsink for engine waste heat, it will still not reach therequired injection temperature without implementation of the vaporiser303 and optionally the heater 304, depending on the chosen meteringtemperature. Further, even in a complex cycle configuration in which theheat of combustion products is recuperated into the hydrogen fuel, ithas been determined that at certain points in the operational envelopethere will be insufficient heat output from the engine to raise the fueltemperature to the injection temperature. Such occasions may include,for example, ground start, in-flight relight, end of cruise idle, etc.

An example configuration of the vaporiser 303 is shown in FIG. 5. Such aconfiguration may also be used for the heater 304.

The vaporiser 303 comprises an offtake 501 from a main fuel conduit 502.The amount of hydrogen bled from the main fuel conduit 502 is controlledby a valve 503. In operation, of the order of around 1 percent of thehydrogen fuel flow through the main fuel conduit 502 is bled for use inthe vaporiser 303.

As described previously, hydrogen has very high specific and latent heatcapacities; however as a gas it has a very low molecular weight anddensity, and thus it can be challenging to exchange heat in a compactway. Thus the vaporiser 303 vaporises the hydrogen fuel in the main fuelconduit 502 by combustion of the bled fuel in a burner 504 located inheat exchange relationship with the main fuel conduit 502. In thepresent embodiment, the burner 504 is concentric around the main fuelconduit 502, although it will be appreciated that other arrangements arepossible.

In the present embodiment, air for combustion with the bled hydrogenfuel is bled from the high-pressure compressor 204. Alternatively, itmay be bled from the low-pressure compressor 202. It will be appreciatedthat the air for combustion could be obtained from any other suitablelocation.

In the present example, the air and the bled hydrogen fuel are mixed ina premixer 505, although in alternative embodiments it may be directlyco-injected into the burner with the hydrogen fuel instead. Combustionproducts from the burner 504 are, in an embodiment, exhausted into thebypass duct of the turbofan engine 103. Alternatively, they may beexhausted outside the nacelle.

It should be understood that, in the present example, the products ofcombustion from the burner 504 are not mixed with the fuel in the mainfuel conduit 502. In this respect, the vaporiser 303 therefore differsfrom a pre-burner system as used in staged combustion cycle rocketengines.

In steady state, there is enough heat emanating from the burner 504 toensure vaporisation of the small amount of bled hydrogen fuel. At enginestart or other cold conditions for example, the vaporiser 303 comprisesa preheater 506 to ensure vaporisation of the bled hydrogen fuel priorto mixing with air in the premixer 505. In a specific embodiment, thepreheater 506 comprises an electric heating element, for example a coil.Alternatively, the preheater 506 could be simply configured as a boilvolume, in which the ambient conditions therein contain sufficiententhalpy to boil the initial flow of bled hydrogen fuel prior todelivery to the premixer 505 and the burner 504.

Embodiments of the metering device 302 are illustrated in FIGS. 6A and6B.

Fuel flow on a conventional liquid-fuelled aero engine is typicallycontrolled by means of a pressure regulating valve and a profiledtranslating spill valve which returns a proportion of the flow suppliedby the pump back to the pump inlet. However, because hydrogen has anextremely low density and viscosity, it has a strong tendency to leakthrough any gap. A control system that relies on close clearances tominimise leakages will be highly problematic with hydrogen as the fuel,since there will be significant leakage with even very tight clearancesand the significant thermal variations in a hydrogen system willpreclude very tight clearances.

In the present embodiments, therefore, the metering device 302 uses afixed orifice which inherently has no moving parts and may therefore besealed.

A first embodiment of the metering device 302 is shown in FIG. 6 andcomprises a choked sonic orifice 601 located in the main fuel conduit302. Thus, in operation, the flow is through the orifice is choked, i.e.it has a Mach number of 1. The flow is therefore a function only of thearea of the orifice and upstream pressure and temperature, measured inthis embodiment by a sensor 602. In order to ensure the orifice remainschoked, the orifice 601 comprises an exit with no expansion, i.e. it issharp-edged, and the ratio of upstream to downstream pressures is set tobe at least the critical pressure ratio which, for hydrogen (a diatomicgas) is around 1.9.

Flow control is then achieved simply by adjusting the upstream pressuredelivered by the pump 301, the upstream temperature being measured andthe orifice area being known.

As an alternative, the metering device 302 could comprise a fixed butunchoked orifice across which a pressure differential may be measuredacross upstream and downstream taps using an appropriate sensor. Massflow may then be derived with knowledge of upstream and downstreampressures and temperatures and the geometry of the fixed orifice.

Complex Cycles

As described previously, it is envisaged that the fuel delivery system201 and fuel injection system 206 may be used in an embodiment of thecore gas turbine 105 implementing a simple cycle as described withreference to FIG. 2, possibly with fuel cooling of engine or gearbox oilor cooling air. Alternatively, the core gas turbine engine 105 mayimplement a complex cycle.

A first embodiment of such a complex cycle is shown in FIG. 7 with likereference numerals used for matching features. In this example, theturbofan engine 103 and core gas turbine 105 are unchanged from theirarrangement in FIG. 2, save for the addition of a recuperator 701located between the low-pressure turbine 209 and core nozzle 210. Therecuperator 701 forms part of the fuel heating system and is operable toheat hydrogen fuel by the exhaust stream of the core gas turbine 105. Inthis way, less fuel may be required to heat the hydrogen fuel to theinjection temperature, increasing cycle efficiency.

In an embodiment, the recuperator 701 is a spiral-wound recuperator,which reduces the likelihood of fracture due to thermal expansion andcontraction.

Another embodiment of a complex cycle is shown in FIG. 8, which buildson the cycle of FIG. 7 with the inclusion of a fuel turbine 801. It willbe appreciated that substantial energy recovery may be achieved from theexhaust stream if it is accepted that less thrust will be developed bythe core nozzle 210. Thus, it is possible to heat the hydrogen fuelbeyond the required fuel injection temperature and to recover work inthe fuel turbine 801, which may be used to drive a load 802. In thisexample the load 802 is an electrical generator. In a specificembodiment, the electrical generator powers the fuel pump 301.Alternatively, the load could be the second shaft 212, with anappropriate drive mechanism being provided. In this way, the fuelturbine 801 augments the low-pressure turbine 209. It will beappreciated that other engine loads such as oil pumps etc. could also bedriven by the fuel turbine 801.

Additionally or alternatively, as shown in FIG. 9 it is possible performfurther recuperation by using the hydrogen fuel to cool the combustor207. Gas turbine combustors feature a liner needs to be cooled tomaintain its mechanical integrity.

In conventional liquid-fuelled aero engines the combustor liner iscooled by the airflow drawn from atmosphere and which has passed throughthe compression system. This is typically via a single pass system inwhich the air passes through holes in the liner and to enter the mainheat release region. Hence this air cannot be part of the combustionprocess and therefore leads to an increase in emissions and a decreasein cycle efficiency.

Thus, in an embodiment, the hydrogen fuel is flowed around the liner ofthe combustor 207. This scheme may be achieved by provision of forexample helical cooling channels around the combustor 207 through whichthe hydrogen fuel may flow prior to injection.

Additionally or alternatively, as shown in FIG. 10 it is possible toprovide intercooling and twin-pass recuperation.

In this embodiment, an intercooler 1001 is provided in the interstageduct 203 between the low-pressure compressor 202 and the high-pressurecompressor 204 for cooling low-pressure compressor discharge air by thehydrogen fuel. In this way, the amount of compression work required tobe performed by the high-pressure compressor 204 is reduced.

In this specific embodiment, a second recuperator 1002 is providedbetween the low-pressure turbine 209 and the recuperator 701 for furtherrecuperative heating of the hydrogen fuel.

Thus, in this example, hydrogen fuel is first heated by the recuperator701 to a temperature less than the low-pressure compressor 202 dischargeair, which heats it further in the intercooler 1001. Further heatingoccurs in the second recuperator 1002, which has an inlet temperaturehigher than the recuperator 701. In this way, the temperature differencebetween the hydrogen fuel and the core gas turbine exhaust temperatureis maximised in each recuperator.

Additionally or alternatively, as shown in FIG. 11 a sequentialcombustion arrangement may be implemented to facilitate inter-turbinereheat. It will be appreciated that reheat of this type comprisesadditional stages of combustion to raise temperatures back to a maximumcycle temperature after a first stage of expansion. Along withintercooling, this moves the overall engine cycle closer to an Ericssoncycle, improving thermal efficiency substantially. In this specificexample, the high-pressure turbine 208 is a multi-stage turbine and areheat fuel injection system 1101 and reheat combustor 1102 arestationed between two of the stages 208A and 208B of the high-pressureturbine 208. Alternatively, the reheat fuel injection system 1101 andreheat combustor 1102 may be stationed between the high-pressure turbine208 and the low-pressure turbine 209.

Direct Fuel Injection System

Due to its wide flammability limits and reaction rates, there issignificant risk of flashback in hydrogen fuel injection systems. Thusit is preferable to utilise the direct injection principle with lowmixing times and high velocities, as opposed to attempting any form ofpremixing. In order to minimise formation of oxides of nitrogen,residence time at high temperate must also be minimised. Theseconstraints therefore favour a miniaturisation of the individual fuelinjectors, sometimes referred to as “micromix” injectors.

FIGS. 12A and 12B illustrate two possible arrangements of the fuelinjection system 206. It will be appreciated that in the presentembodiment the core gas turbine 105 employs an annular combustionsystem, and it will be clear how the principles disclosed herein may beadapted e.g. for tubular systems.

In the embodiment of FIG. 12A, the fuel injection system 206 comprises afull annulus 1201 of fuel injector blocks 1202. In the embodiment ofFIG. 12B, the fuel injection system 206 comprises a plurality of sectors1203 each comprising a subset of the totality of fuel injector blocks1202. In both embodiments, the fuel injector blocks 1202 are configuredwith a geometry that substantially tessellates. It will be appreciatedthat the embodiment of FIG. 12A will produce a substantially moreuniform circumferential heat-release profile, reducing the danger of hotstreaks in the combustor 207 and uneven loading of the high-pressureturbine 208, improving performance by reducing cooling requirements.

It is contemplated that the fuel injection system 206 would comprisemany hundreds or even thousands of fuel injector blocks 1202. Forexample, in an embodiment there are from 500 to 2000 fuel injectorblocks, for example 1000 fuel injector blocks 1202.

A first configuration for the fuel injector blocks 1202 will bedescribed with reference to FIGS. 13A to 15B. A second configuration forthe fuel injector blocks 1202 will be described with reference to FIGS.16 to 17B. A third configuration for the fuel injector blocks 1202 willbe described with reference to FIGS. 18A to 20B.

The first configuration for the fuel injector blocks is shown in FIGS.13A and 13B, and will hereinafter be referred to as a rim injector block1301. A first embodiment of the rim injector block 1301 is shown in FIG.13A, and has a quadrilateral, specifically a square, outer profile inthe plane of tessellation. Another embodiment is shown in FIG. 13B, andhas a hexagonal, specifically a regular hexagon, outer profile in theplane of tessellation. It will be appreciated that other outer profilesthat tesselate could be used. In this example, the rim injector block1301 comprises an air admission duct 1302 and a fuel admission aperture1303.

Referring now to FIG. 14A, which is a cross sectional view on I-I ofFIG. 13A, the air admission duct 1302 has an inlet 1401 for receivingair A from the diffuser 205 and an outlet 1402 for delivering air into amixing zone in the combustor 207. The air admission duct 1302 has acentral axis C extending from the inlet 1401 to the outlet 1402. Thefuel admission aperture 1303 is located around the periphery of theoutlet 1402 and is configured to inject hydrogen onto a jet shear layerformed at the outlet 1302 for mixing in the mixing zone. In theembodiment of FIG. 14A, the fuel admission aperture 1303 is configuredto inject hydrogen fuel parallel to the central axis C, as shown byarrows F. FIG. 14B shows the equivalence ratios downstream of the riminjector block 1301, and was obtained by a periodic isothermal CFDsimulation on this configuration. The air admission duct 1302 was sizedwith a 6 millimetre diameter. In this example, a uniform equivalenceratio U was achieved within 80 millimetres of the injection point.

An alternative configuration of the rim injector block 1301 is shown inFIG. 15A, in which the fuel admission aperture 1303 is configured toinject hydrogen fuel perpendicular to the central axis C, as shown byarrows F. FIG. 15B shows the results of a periodic isothermal CFDsimulation on this configuration. Again, the air admission duct 1302 hada 6 millimetre diameter and a uniform equivalence ratio U was achievedwithin 80 millimetres of the injection point.

In both configurations, the injection of fuel onto the jet shear layerfrom the fuel admission aperture 1303 minimises flammable mixtures atvelocities lower than the turbulent flame speed close to the injector.This reduces the risk of flashback.

A second configuration for the fuel injector blocks is shown in FIG. 16,and will hereinafter be referred to as a converging jet injector block1601. The converging jet injector block 1601 comprises a fuel admissionduct 1602 and a plurality of air admission ducts 1603 distributed aroundthe air admission duct. In this specific embodiment, the air admissionducts 1603 are equidistant from the fuel admission duct 1602. Asdescribed previously, the outer profile of the injector block may beconfigured such that multiple blocks substantially tesselate adjacent toone another.

Referring now to FIG. 17A, which is a cross-sectional view on II-II ofFIG. 16, the fuel admission duct 1602 has a central axis C and each airadmission duct 1603 has its own respective axis R defined between eachduct's respective inlet and outlet. In the present embodiment, the fueladmission duct 1602 and air admission ducts 1603 are configured torespectively admit fuel and air without swirl. The respective axis R ofeach air admission duct 1603, is inclined towards the central axis C ofthe fuel admission duct 1602, such that emerging fuel F converges on theair A in the mixing zone.

FIG. 17B shows the equivalence ratios downstream of the converging jetinjector block 1601, and was obtained by a periodic isothermal CFDsimulation on this configuration. In this example, the fuel admissionduct 1602 was sized with a 0.5 millimetre diameter and the air admissionducts 1603 were sized with a 0.8 millimetre diameter. A uniformequivalence ratio U was achieved within 30 millimetres of the injectionpoint.

A third configuration for the fuel injector blocks is shown in FIGS. 18Aand 18B, and will hereinafter be referred to as a jet matrix injectorblock 1801. The converging jet injector block 1801 comprises a fueladmission duct 1802 and a plurality of air admission ducts 1803distributed around the air admission duct. In the present embodiment,the fuel admission duct 1802 and air admission ducts 1803 are configuredto respectively admit fuel and air without swirl.

In FIG. 18A, the outer profile of the jet matrix injector block 1801 isa quadrilateral and in FIG. 18B it is a hexagon.

In an embodiment, the jet matrix injector block 1801 comprises a number2N of air admission ducts 1803 that meet the following criteria: 2N iseven and equal to 6 or more, i.e. 2N≥6. Thus for example in FIG. 18A itcan be seen that N=4 and thus there are 8 air admission ducts 1803,whilst in FIG. 18A it can be seen that N=6 and thus there are 12 airadmission ducts 1803.

The air admission ducts 1803 are then distributed around the fueladmission duct 1802 such that they lie on the periphery of an N-gon(i.e. a polygon with N sides) centred on the central axis C of the fueladmission duct 1802. In a specific embodiment, N of the air admissionducts 1803 are arranged at a respective vertex of the N-gon, and theother N of the air admission ducts 1803 are arranged on a respectiveedge of the N-gon. For example, the other N of the air admission ducts1803 may be arranged at the midpoint of their respective edge.

Taking FIG. 18A as a worked example, N=4, therefore there are 2N=8 airadmission ducts 1803. These are located on the periphery of a 4-gon,i.e. a square. Four of the air admission ducts 1803 are located at thevertices (i.e. the corners) of the square, and the other four arelocated on a respective edge of the square. In this example they arelocated at a midpoint of the edges.

FIG. 19A is a cross-sectional view on III-III of FIGS. 18A and 18B, inwhich it may be seen that in this embodiment the air admission ducts1803 are configured to lie parallel to the fuel admission duct 1802.Thus, each air admission duct 1803 has a respective axis R definedbetween each duct's inlet and outlet, whilst the fuel admission duct1802 has a central axis C defined between its inlet and outlet. Therespective axes R of the air admission ducts 1803 are parallel with thecentral axis C.

FIG. 19B shows the equivalence ratios downstream of the jet matrixinjector block 1901, and was obtained by a periodic isothermal CFDsimulation on this configuration. In this example, the fuel admissionduct 1802 was sized with a 0.5 millimetre diameter and the air admissionducts 1803 were sized with a 0.8 millimetre diameter. A uniformequivalence ratio U was achieved within 10 millimetres of the injectionpoint.

By defining fuel injector blocks 1202 of small scale relative to theoverall size of the fuel injection system annulus, the flow field in thecombustor 207 becomes self-similar and substantially invariant overdifferent practical sizes. An example is shown in FIGS. 20A and 20B ofhow the core gas turbine 105 may undergo a power scaling, i.e. the useof a substantially common design for two different power levels. In thisexample, the fuel injection system 206 is sized for an engine with powerP in FIG. 20A, and a power 2P in FIG. 20B. However, it will be seen thatthe size of the fuel injector blocks 1202 has not changed between thetwo designs, there has simply been an increase of the number making upthe overall annulus.

Thus, in an industrial setting, the design process for a newspecification engine may simply comprise obtaining a design of astandardised fuel injector block, such as blocks 1301, 1601 and 1801.The standard specification for such an injector block would comprise itscapability in terms of fuel mass flow performance and its dimensions andgeometry. Engine performance data, typically derived prior to detailedcomponent design, would set the required fuel mass flow requirements forthe new engine type.

A simple evaluation of the quantity of standardised fuel injector blocksthat meets the fuel mass flow requirements for the engine may then beperformed. This would not require any dimensional scaling of thestandardised fuel injectors, and indeed this would be discouraged as theflow field would change.

Referring again briefly to FIG. 2, it can be seen that the adoption of amicromix-type direct fuel injection system as described hereinfacilitates a much shorter combustor 207. Thus for the same overallaxial length of the core gas turbine 105, it is possible to increase theaxial length available for the diffuser 205. In this way, improvedcontrol over the diffusion process may be achieved rather than theconventional dump diffuser designs utilised on conventionalliquid-fuelled aero engines which require much longer combustor lengthsdue to the slower combustion processes associated with liquidhydrocarbon fuels. Thus in an embodiment, the axial length of thediffuser 205 is greater than the axial length of the combustor 207. Inan embodiment, the diffuser is a faired diffuser which improves theuniformity of the flow delivered to the fuel injection system 206.

Various examples have been described, each of which comprise variouscombinations of features. It will be appreciated by those skilled in theart that, except where clearly mutually exclusive, any of the featuresmay be employed separately or in combination with any other features andthus the disclosed subject-matter extends to and includes all suchcombinations and sub-combinations of the or more features describedherein.

1. A complex cycle gas turbine engine for an aircraft with hydrogen fuelsupply, comprising, in fluid flow series, a core gas turbine and arecuperator for heating hydrogen fuel prior to combustion by turbineexhaust products.
 2. The gas turbine engine of claim 1, in which thecore gas turbine engine comprises: a low-pressure spool having alow-pressure compressor driven by a low-pressure turbine; ahigh-pressure spool having a high-pressure compressor driven by a lowhigh-pressure turbine.
 3. The gas turbine engine of claim 2, furthercomprising a fan driven by the low-pressure spool, optionally via areduction gearbox.
 4. The gas turbine engine of claim 2, in which thecore gas turbine comprises a combustor between the high-pressurecompressor and the high-pressure turbine, the combustor having a linercooled in a regenerative configuration by the hydrogen fuel.
 5. The gasturbine engine claim 1, further comprising a fuel turbine for receivingheated hydrogen fuel from the recuperator to drive a load.
 6. The gasturbine engine of claim 5, in which the load is: an electricalgenerator; a fuel pump for pumping the hydrogen fuel; a low-pressurespool or a high-pressure spool in the core gas turbine.
 7. The gasturbine engine claim 2, further comprising an intercooler between thelow-pressure compressor and the high-pressure compressor for coolinglow-pressure compressor discharge air by the hydrogen fuel.
 8. The gasturbine engine of claim 7, further comprising a second recuperator forfurther heating hydrogen fuel received from the intercooler.
 9. The gasturbine engine of claim 8, in which the second recuperator is stationeddownstream of the recuperator.
 10. The gas turbine engine of claim 1,further comprising a reheat combustor.
 11. The gas turbine engine ofclaim 10, in which the core gas turbine engine comprises a multi-stageturbine and the reheat combustor is stationed between two of the stageson the turbine.
 12. The gas turbine engine of claim 2, furthercomprising a reheat combustor stationed between the high-pressureturbine and the low-pressure turbine.
 13. The gas turbine engine claim1, further comprising a fuel delivery system for delivering the hydrogenfuel from a cryogenic storage system to the recuperator, the fueldelivery system including a pump, a metering device, and a fuel heatingsystem for heating the hydrogen fuel.
 14. The gas turbine engine ofclaim 13, in which the fuel heating system comprises a vaporiserconfigured to vaporise liquid hydrogen from the cryogenic storagesystem.
 15. The gas turbine engine of claim 14, in which the vaporisercomprises a fuel offtake for diverting a portion of the hydrogen fuelfrom a fuel conduit for combustion in a burner located in heat exchangerelationship with the fuel conduit.
 16. The gas turbine engine of claim16, in which the vaporiser comprises a boil volume or an electricheating element for initial heating of liquid hydrogen if no vaporisedhydrogen fuel is available.
 17. The gas turbine engine of claim 13, inwhich the metering device is a fixed orifice and flow rate is controlledby varying the pressure ratio across the orifice.
 18. The gas turbineengine of claim 17, in which the metering device comprises a sonic fixedorifice configured to operate in a choked condition, and flow rate iscontrolled by varying pressure upstream of the sonic fixed orifice. 19.The gas turbine engine of claim 13, in which the fuel heating systemcomprises one or more heat exchangers for heating the hydrogen fuel byheat from the core gas turbine.
 20. The gas turbine engine of claim 13,in which one or more heat exchangers are oil-fuel heat exchangers forcooling engine oil or gearbox oil by the hydrogen fuel.